Turbomachine thermal management

ABSTRACT

An example turbomachine assembly includes, among other things, a nose cone of a turbomachine. The nose cone provides an aperture that communicates air to an interior of the nose cone.

BACKGROUND

This disclosure relates to thermal management within a turbomachine.

Turbomachines, such as gas turbine engines, typically include a fansection, a compressor section, a combustor section, and a turbinesection. These sections are axially sequentially arranged. Turbomachinesmay employ a geared architecture connecting the fan section and thecompressor section.

In many turbomachines, some of the air in the compressor section isreturned to the fan section. This returned air helps limit ice build-up,buffers seals, etc.

SUMMARY

A turbomachine assembly according to an exemplary aspect of the presentdisclosure includes, among other things, a nose cone of a turbomachine,the nose cone providing an aperture that communicates air to an interiorof the nose cone.

In a further non-limiting embodiment of the foregoing turbomachineassembly, the aperture may be provided by a circumferential slot in thenose cone.

In a further non-limiting embodiment of either of the foregoingturbomachine assemblies, the nose cone may establish a portion of aradially inner boundary of flow entering a fan section of theturbomachine.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the nose cone and the aperture may be coaxiallyaligned at a rotational axis of the turbomachine.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the assembly may include a pump thatcommunicates air through the nose cone. The pump may be at leastpartially housed within the nose cone.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the pump may be rotated to communicate airthrough the aperture.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the pump may be driven by a low shaft to moveair through the aperture.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the pump may be an impeller.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the assembly may include a heat exchanger thatreceives air communicated through the aperture.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the heat exchanger may be forward of a fanblade platform of the turbomachine relative to a direction of flowthrough the turbomachine.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, air that has been communicated through theaperture may carry thermal energy from the heat exchanger.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, air that has moved through the aperture may addthermal energy to the nose cone.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, air that has moved through the aperture may addthermal energy to one of the fan blade platforms, front center bodystruts, or low pressure compressor inlet guide vane airfoils.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, air that has moved through the aperture maybuffer a seal.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the turbomachine may have core and bypassflowpaths. Air that has moved through the aperture may be communicatedfrom an inner diameter of the core flowpath to an outer diameter of thecore flowpath.

A turbomachine assembly according to another exemplary aspect of thepresent disclosure includes, among other things, a nose cone having aninterior extending forward a fan section of a turbomachine, and a pumpthat communicates air to the interior.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies, the pump may pressurize the air that iscommunicated to the interior.

A method of communicating air within a turbomachine according to anotherexemplary aspect of the present disclosure includes, among other things,moving air along a path from a first position outside a nose coneinterior directly to a second position within the nose cone interior.The path is forward an array of fan blades within a turbomachine.

In a further non-limiting embodiment of the foregoing method ofcommunicating air within a turbomachine, the method may include heatingthe nose cone using air that has moved along the path.

In a further non-limiting embodiment of either of the foregoing methodsof communicating air within a turbomachine, the method may includebuffering of bearing compartment seals using air that has moved alongthe path. In a further non-limiting embodiment of any of the foregoingmethods of communicating air within a turbomachine, the air at thesecond position may be pressurized relative to the air at the firstposition.

In a further non-limiting embodiment of either of the foregoing methodsof communicating air within a turbomachine, the method may includelimiting flow that has moved along the path from leaking into a coreflowpath of the turbomachine using a seal.

DESCRIPTION OF THE FIGURES

The various features and advantages of the disclosed examples willbecome apparent to those skilled in the art from the detaileddescription. The figures that accompany the detailed description can bebriefly described as follows:

FIG. 1 shows a partial section view of an example turbomachine.

FIG. 2 shows an upper axial half of a front end of the FIG. 1turbomachine.

FIG. 3 shows an upper axial half of a front end of another exampleturbomachine.

FIG. 4 shows an upper axial half of a front end of yet another exampleturbomachine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example turbomachine, which is a gasturbine engine 20 in this example. The gas turbine engine 20 is atwo-spool turbofan gas turbine engine that generally includes a fansection 22, a compressor section 24, a combustion section 26, and aturbine section 28. Other examples may include an augmentor section (notshown) among other systems or features.

Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with turbofans. Thatis, the teachings may be applied to other types of turbomachines andturbine engines including but not limited to one or three-spoolarchitectures. The number of stages in each engine module also may vary.

In the example engine 20, the fan section 22 drives air along a bypassflowpath B while the compressor section 24 drives air along a coreflowpath C. Compressed air from the compressor section 24 is mixed withfuel and ignited in the combustion section 26. The products ofcombustion expand through and drive the turbine section 28.

The example engine 20 generally includes a low-speed spool 30 and ahigh-speed spool 32 mounted for rotation about an engine longitudinalaxis A relative to an engine static structure 36. The low-speed spool 30and the high-speed spool 32 are rotatably supported by several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively, or additionally, be provided.

The low-speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan rotor 42, a low-pressure compressor 44, and alow-pressure turbine 46. As illustrated, the inner shaft 40 is connectedto the fan rotor 42 through a speed reduction device 48 to drive the fanrotor 42 at a lower speed than the low-speed spool 30. The low-speedspool 30 is low-speed relative to the rotational speed of the high-speedspool 32. The high-speed spool 32 includes an outer shaft 50 thatcouples a high-pressure compressor 52 and high-pressure turbine 54.

The combustion section 26 includes a circumferential array of fuelnozzles 56 generally arranged axially between the high-pressurecompressor 52 and the high-pressure turbine 54.

An optional mid-turbine frame 58 of the engine static structure 36 isgenerally arranged axially between the high-pressure turbine 54 and thelow-pressure turbine 46. The mid-turbine frame 58 supports the bearingsystems 38 in the turbine section 28 in this particular configuration.

The inner shaft 40 and the outer shaft 50 are concentric and rotate viathe bearing systems 38 about the engine central longitudinal axis A,which is collinear with the longitudinal axes of the inner shaft 40 andthe outer shaft 50. The shafts are counter-rotating, but may beco-rotating in another example.

In the example engine 20, the airflow moved along the core flowpath C iscompressed by the low-pressure compressor 44 then the high-pressurecompressor 52, mixed and burned with fuel in the combustors fuel nozzles56, then expanded over the high-pressure turbine 54 and low-pressureturbine 46. The high-pressure turbine 54 and the low-pressure turbine 46rotatably drive the respective high-speed spool 32 and low-speed spool30 in response to the expansion.

In some non-limiting examples, the engine 20 is a high-bypass gearedaircraft engine. In a further example, the engine 20 bypass ratio isgreater than about 6 (6 to 1).

The speed reduction device 48 of the example engine 20 includes anepicyclic gear train, such as a planetary, star, or differential gearsystem. The example epicyclic gear train has a gear reduction ratio ofgreater than about 2.2 (2.2 to 1).

The low-pressure turbine 46 pressure ratio is pressure measured prior toinlet of low-pressure turbine 46 as related to the pressure at theoutlet of the low-pressure turbine 46 prior to an exhaust nozzle of theengine 20. In one non-limiting embodiment, the bypass ratio of theengine 20 is greater than about 10 (10 to 1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low-pressure turbine 46 has a pressure ratio that is greater thanabout 5 (5 to 1). The speed reduction device 48 of this embodiment is anepicyclic gear train with a gear reduction ratio of greater than about2.4 (2.4 to 1). It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a geared architectureengine and that the present disclosure is applicable to other gasturbine engines including direct drive turbofans.

In this embodiment of the example engine 20, a significant amount ofthrust is provided by the flow through the bypass flowpath B due to thehigh bypass ratio. The fan section 22 of the engine 20 is designed for aparticular flight condition—typically cruise at about 0.8 Mach and about35,000 feet (10,668 meters). This flight condition, with the engine 20at its best fuel consumption, is also known as bucket cruise. ThrustSpecific Fuel Consumption (TSFC) is an industry standard parameter offuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22, typically without the use of a Fan Exit Guide Vane system.The low Fan Pressure Ratio according to one non-limiting embodiment ofthe example engine 20 is less than 1.45.

Low Corrected Fan Tip Speed is the actual fan tip speed divided by anindustry standard temperature correction of “T”/518.7^(0.5). Trepresents the ambient temperature in degrees Rankine. The Low CorrectedFan Tip Speed according to one non-limiting embodiment of the exampleengine 20 is less than about 1150 fps (351 m/s).

Referring now to FIG. 2, the fan rotor 42 of the gas turbine engine 20includes a plurality of fan blades 62 extending radially from the axisA. The example gas turbine engine 20 also includes a nose cone 66. Inthis example, the nose cone 66 extends axially forward from the fanblades 62. This nose cone 66 is sometimes referred to as a spinner.

The nose cone 66, in this example, establishes the forwardmost portionof the gas turbine engine 20. The nose cone 66 directs and streamlinesflow entering the gas turbine engine 20. An outer surface 70 of the nosecone 66 establishes a portion of a radially inner boundary of a flowpathfor flow entering the engine 20. The nose cone 66 may include one pieceor several individual pieces.

The nose cone 66 has an interior 72. The nose cone 66 provides anaperture 74 that communicates flow to the interior 72. The aperture 74may be positioned at a front 76 of the nose cone 66 as shown, or in someother position aft the front 76 of the nose cone 66. The exampleaperture 74 and the nose cone 66 are coaxially aligned at the axis A.Other alignments of the aperture 74 are possible. A line-of-siteblockage structure may be incorporated to protect the system frombirdstrike damage.

In this example, a pump, such as an impeller 80, is used to increase theamount of air entering the interior. The impeller 80 is positioned atleast partially within the interior 72 of the nose cone 66. The impeller80 is driven by rotation during operation. Other example pumps mayinclude an axial fan.

Notably, some examples may not utilize the impeller 80 or any otherpump. In such examples, air enters the interior 72 through the aperture74 due to forward movement of the engine 20. As appreciated, designsincorporating pumps typically move more air to the interior 72 thandesigns that do not incorporate any pumps. Air can enter the interior 72through aperture 74 when the exit pressure at the bypass duct innerdiameter discharge location can be suppressed below that of the inlet tothe nose cone 66. This can occur when the engine 20 is not movingforward, but still turning, like at ground idle.

During operation, the example impeller 80 is rotated by a shaft 82,which is rotated by a shaft 88. The example shaft 88 is a flex couplingshaft. The shaft 88 is driven by the inner shaft 40 of the low-speedspool 30 (FIG. 1). Rotating the shaft 88 rotates a ring gear 96 of thegeared architecture 48, which then rotates a gear 98 to rotate the shaft82.

The example engine 20 includes a heat exchanger 100 positioned near thenose cone 66. In this example, at least a portion of the heat exchanger100 is forward of the fan blades 62. Air that has entered the interior72 through the aperture 74 is moved through the heat exchanger 100 toprovide a cold side fluid.

The example heat exchanger 100 is a fin and tube type heat exchanger100, although other type of heat exchangers may also be used. The heatexchanger 100 is annular and distributed about the axis A. In anotherexample, several separate heat exchangers are distributed about the axisA. A support 102 extending axially from a stationary carrier 90 of thegeared architecture 48 supports the heat exchanger 100 in this example.The support 102 may also support conduits (not shown) carrying lubricantto the heat exchanger 100.

Air that has entered the interior 72 of the nose cone 66 is pressurizedrelative to the air outside the nose cone 66. In one example, thedifference in pressure between the air within the interior 72 of thenose cone 66 and the air outside the nose cone 66 ranges from 0.5 to 10psi (3.45 kPa to 68.95 kPa).

Air that has entered the interior 72 of the nose cone 66 is raised to ahigher temperature relative to the air outside the nose cone 66. In oneexample, the difference in temperature between the air within theinterior 72 of the nose cone 66 and the air outside of the nose cone 66ranges from 100° F. to 200° F. (37.78° C. to 93.33° C.).

Lubricant, such as oil from the engine 20, provides the warm side fluidfor the heat exchanger 100. Air that has entered the interior 72 ismoved through the heat exchanger 100 to remove thermal energy from thelubricant.

The lubricant may be a mixture of lubricant from several areas of thegas turbine engine 20, or may be lubricant from a subset of alubrication system of the engine 20, such as a fan drive gear systemlubrication system. Although the air that has entered the interior 72 isincreased in temperature relative to the air outside the nose cone 66,the air that has entered the interior 72 is still significantly coolerthan the lubricant within the heat exchanger 100 and therefore can stillremove heat from it.

In this example, the heat exchanger 100 is configured to rejection atleast 2,900 BTU/min at idle and 5,000 BTU/min and take-off.

Some of the air that has moved through the heat exchanger 100 movesalong path P₁ and contacts an interior facing side of the nose cone 66.This air adds thermal energy to the nose cone 66, which inhibits iceformation. As appreciated, air moving along path P₁ includes morethermal energy that the air moving into the heat exchanger 100 becauseair that has moved through the heat exchanger 100 has absorbed thermalenergy from the lubricant moving through the heat exchanger 100.

Some of the air that has moved through the heat exchanger 100 movesalong path P₂ and buffers a carbon seal 106 within bearings 38 of thegas turbine engine 20. The carbon seal 106 is a forward carbon seal ofthe front bearing compartment of the engine 20 in this particular case.The seal 106, in this example, is biased axially rearward by the air.

Most, if not all, of the air that has moved through the heat exchanger100 eventually moves along path P₃. A portion of the path P₃ extendsradially inward the fan blade platform 108. If additional flow area isneed to permit the required flow rate, additional apertures may beincluded through the Fan Blade Hub Web, which is the same part as thefan rotor 42 in this example, just at a radially inboard. The path P₃also extends through an inlet guide vane 110 and a front center bodystrut 112. The inlet guide vane 110 and the front center body 112include internal passages and extend radially across the core flowpath Cof the engine 20. The path P₃ terminates at the bypass flowpath B of thegas turbine engine 20 behind the FEGV (fan exit guide vanes). Thus, airthat has entered the interior 72 through the aperture 74 is eventuallyexpelled into the bypass flowpath B at an axial location having a staticpressure that is low enough to enable sufficient flow through the heatexchanger 100.

In addition to the nose cone 66, the air that has moved through the heatexchanger 100 also adds thermal energy to inhibit ice formation on otherstructures, such as the fan blade platform 108, the inlet guide vane110, and the front center body struts 112.

In this example, a seal structure 120 seals the interface between thefan rotor 42 and a stationary structure of the gas turbine engine 20.The seal structure 120 limits flow escaping from the path P₃, orinterior 72, to a position directly aft the fan blade 62. Such leakageinto this area may undesirably introduce turbulence to the flow enteringthe core flowpath C of the gas turbine engine 20.

Referring to FIG. 3, in another example engine 20 a, the impeller 80 isselectively driven by an electric motor 122, which allows relativelyinfinite adjustments of the rotational speed of the impeller 80. Themotor 122 may be used to rotate the impeller 80 when the engine 20 is onthe ground, at idle, or at the top of descent. When the motor 122 is notin use, the impeller 80 may be driven using the shaft 88.

In this example, a clutch 126 is moved between an engaged position and adisengaged position to selectively drive the impeller 80 using theelectric motor 122 or using the shaft 88. The clutch 126 is an AirTurbine Starter (ATS)-type ratchet clutch in this example.

In the engaged position, the clutch 126 couples together the shaft 128to a shaft 82 a such that rotation of the shaft 128 rotates the shaft 82a. The shaft 82 a is directly connected to the impeller 80. That is,rotating the shaft 82 a rotates the impeller 80.

The shaft 128 is rotated by the ring gear 96 of the geared architecture48, which is rotated by the shaft 88. The motor 122 is not driving theshaft 82 a when the clutch 126 is engaged. When the motor 122 isrunning, and driving the impeller 80, the clutch 126 disengages suchthat the shaft 82 a is free to be rotated by the electric motor 122relative to the shaft 128. In FIG. 3, the clutch 126 is shown in adisengaged position, which corresponds to the electric motor 122 drivingthe impeller 80.

Referring to FIG. 4, in yet another example engine 20 b, the impeller isselectively rotated by the motor 122 or the shaft 132, again selected bya clutch 126. However, in this example, a geared architecture 140 isused to alter the rotational speed of the shaft 132 relative to theshaft 88. In this example, the shaft 88 rotates the ring gear 96 torotate the geared architecture 140. The geared architecture 140, inturn, rotates the shaft 132 which rotates a shaft 82 b that rotates theimpeller 80 when the clutch 126 is in the engaged position.

In this example, the stationary carrier 90 of the geared architecture 48includes extensions 144 that support gears 148 of the gearedarchitecture 140. The gears 148 are planetary gears in this example.Rotating the geared architecture 140 rotates the shaft 82 to drive theimpeller 80. The geared architecture 140, in this example, has a gearratio of about five (5 to 1). That is, when the shaft 88 is used torotate the impeller 80, one rotation of the shaft 88 rotates the shaft82 (and the impeller 80) five times.

When the impeller 80 is not rotated by the shaft 88, the impeller 80 isdriven by the electric motor 122. The clutch 126 controls the selectiverotation of the impeller 80 using the electric motor 122 or shaft 88.

Utilizing the geared architecture 140 enables the impeller 80 to rotateat a faster speed than the shaft 88 when the clutch 126 is engaged. Inone example, the gear ratio of the geared architecture 140 is about five(5 to 1). That is, one rotation of the shaft 88 rotates the impeller 80five times.

Features of the disclosed examples include a thermal management strategythat provides increased space within a core engine nacelle. Anotherfeature is providing thermal energy to limit ice formation on variouscomponents at a front of an engine without using a dedicated flow ofbleed air from other areas of the engine. Yet another feature isincreasing Line Replaceable Unit (LRU) capability since the heatexchanger is located in the nose cone. LRU refers to components externalto the engine core (pumps, heat exchangers, etc.) that are relativelyeasy to change out if they become damaged. Still another feature isgreater flexibility to adjust nacelle contours around an engine, whichprovides a potential performance benefit.

Engine heat loads are typically handled via a combination of fuel/oiland air/oil coolers in the prior art. And although heat loads aretypically lower at idle and near idle conditions, the associated lowerfuel flow rate at these conditions significantly lessens the ability ofthe fuel/oil coolers to reject heat. The specific features of theembodiments shown in FIGS. 3 and 4, specifically the TMS pump beingrotated by a motor or a dedicated gear set at much higher speeds thancould be accomplished by directly driving the pump with a turbine shaft,thus compensate for the natural drop-off in heat rejection capability inprior art configurations at idle and near-idle conditions.

Rotating the pump via the planetary gear system as shown in the FIG. 4embodiments allows the pump to rotate faster than the shaft rotated bythe turbine. In this manner the disclosed examples effectively replacesthree systems (spinner anti-ice, front bearing compartment buffering,and TMS cooling).

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. Thus, the scope of legal protectiongiven to this disclosure can only be determined by studying thefollowing claims.

We claim:
 1. A turbomachine assembly, comprising: a nose cone of aturbomachine, the nose cone providing an aperture that communicates airto an interior of the nose cone; a plurality of fan blades distributedcircumferentially about a rotational axis of the turbomachine; and aninlet to a low pressure compressor section of the turbomachine, theinlet positioned axially downstream from the plurality of fan bladesrelative to a general direction of flow through the turbomachine.
 2. Theturbomachine assembly of claim 1, wherein the aperture is provided by acircumferential slot in the nose cone relative to a direction of flowthrough the turbomachine.
 3. The turbomachine assembly of claim 1,wherein the nose cone establishes a portion of a radially inner boundaryof flow entering a fan section of the turbomachine.
 4. The turbomachineassembly of claim 1, wherein the nose cone and the aperture arecoaxially aligned at a rotational axis of the turbomachine.
 5. Theturbomachine assembly of claim 1, including a pump that is configured tocommunicate air through the nose cone, the pump at least partiallyhoused within the nose cone.
 6. The turbomachine assembly of claim 5,wherein the pump is configured to rotate to communicate air through theaperture.
 7. The turbomachine assembly of claim 6, wherein the pump isconfigured to be driven by a low shaft to move air through the aperture.8. The turbomachine assembly of claim 5, wherein the pump is animpeller.
 9. The turbomachine assembly of claim 1, including a heatexchanger that is configured to receive air communicated through theaperture.
 10. The turbomachine assembly of claim 9, wherein the heatexchanger is forward of a fan blade platform of the turbomachinerelative to a direction of flow through the turbomachine.
 11. Theturbomachine assembly of claim 9, wherein air that has been communicatedthrough the aperture carries thermal energy from the heat exchanger. 12.The turbomachine assembly of claim 1, wherein air that has moved throughthe aperture adds thermal energy to the nose cone.
 13. The turbomachineassembly of claim 1, wherein air that has moved through the aperture mayadd thermal energy to at least one of a fan blade platform, front centerbody strut, or low pressure compressor inlet guide vane airfoils. 14.The turbomachine assembly of claim 1, wherein air that has moved throughthe aperture is used to buffer a seal.
 15. A turbomachine assembly,comprising: a nose cone of a turbomachine, the nose cone providing anaperture that communicates air to an interior of the nose cone, whereinthe turbomachine has a core flowpath, and air that has moved through theaperture is communicated from an inner diameter of the core flowpath toan outer diameter of the core flowpath.
 16. A turbomachine assembly,comprising: a nose cone having an interior extending forward from a fansection of a turbomachine; a pump that is configured to communicate airto the interior; and an inlet to a low pressure compressor section ofthe turbomachine, the inlet positioned downstream from fan blades of thefan section relative to a general direction of flow through theturbomachine.
 17. The turbomachine assembly of claim 16, wherein thepump is configured to pressurize the air that is communicated to theinterior.
 18. The turbomachine assembly of claim 16, including aline-of-site blocker configured to provide bird strike protection.
 19. Amethod of communicating air within a turbomachine, comprising: movingair along a path from a first position outside a nose cone interiordirectly to a second position within the nose cone interior, wherein thepath is forward an array of fan blades within a turbomachine, whereinthe array of fan blades is forward an inlet to a low pressure compressorsection of the turbomachine.
 20. The method of claim 19, includingheating the nose cone using air that has moved along the path.
 21. Themethod of claim 20, including buffering of bearing compartment sealsusing air that has moved along the path.
 22. The method of claim 19,wherein the air at the second position is pressurized relative to theair at the first position.
 23. The method of claim 19, includinglimiting flow from that has moved along the path from leaking into acore flowpath of the turbomachine using a seal.